This invention relates generally to a combustor assembly for a gas turbine engine. More particularly, this invention relates to an interface between a combustor assembly and a fixed turbine vane portion of a gas turbine engine.
A gas turbine engine typically includes a combustor for igniting a mixture of fuel and compressed air to produce a gas flow. The combustor typically includes an outer shell supporting a plurality of inner heat shields. The inner heat shields are exposed to elevated temperatures produced by ignition of the fuel-air mixture and the resulting gas flow.
Gas flow exiting the combustor enters a fixed array of turbine vanes that directs gas flow to downstream rotating turbine blades. The fixed vanes are intermediate the combustor and the rotating turbine blades. Typically, the support shell and heat shield articles at the aft end of the combustor module terminate at a common axial position or plane upstream of the fixed vanes. The transition of this dual-wall combustor liner system to the downstream endwall or platform (inner and outer diameter flow path surfaces of the turbine vane cascade) create a seam, step or interrupted surface between internal surfaces of the combustor and the surfaces at the inner or outer diameter of the fixed vane cascade.
Disadvantageously, such interrupted surfaces at the interface between the fixed vane array and the combustor interfere with cooling and core gas flows exiting the combustor. The insulating layer of cooling air along the inner surface of the combustor is disrupted by the interface with the fixed vane portion causing undesirable mixing of the cooling air with the hot core gases. This can lead to decreases in the cooling effectiveness of the cooling air and promote elevated temperatures or adverse temperature gradients on the combustor and turbine hardware in this region. Additionally, disruption of the gas flow that moves downstream into the fixed vane causes undesirable aerodynamic properties and thermal profiles that can potentially degrade the downstream turbine and, hence, overall engine performance.
Accordingly, it is desirable to develop an interface between a combustor assembly and a turbine assembly that provides a smooth transition of the cooling and core gas flows in vicinity of the exit of the combustor and proximate to the entrance to the downstream turbine vane.